A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
In general, turbine performance and efficiency may be improved by increased combustion gas temperatures. However, increased combustion temperatures can negatively impact the gas turbine engine components, for example, by increasing the likelihood of material failures. Thus, while increased combustion temperatures can be beneficial to turbine performance, some components of the gas turbine engine may require cooling features or reduced exposure to the combustion gases to decrease the negative impacts of the increased temperatures on the components.
Film cooling gas turbine engine components, e.g., by directing a flow of cooler fluid over the surface of the component, can help reduce the negative impacts of elevated combustion temperatures. For example, cooling apertures may be provided throughout a component; the cooling apertures may allow a flow of cooling fluid from within the component to be directed over the outer surface of the component. However, in areas of high curvature of the component, it can be difficult to direct the flow of cooling fluid from the cooling apertures over the outer surface of the component to form a cooling film of fluid. Further, known methods of forming cooling apertures, e.g., by boring or otherwise machining apertures in the component, can be ineffective in producing optimal cooling aperture lengths for controlling bore cooling and in producing cooling apertures having optimal fluid exit surface angles. In addition, known methods of machining cooling apertures are prone to through-hole scarfing and often present challenges to properly positioning the cooling apertures.
Therefore, improved cooling features for gas turbine components that overcome one or more disadvantages of existing cooling features would be desirable. In particular, an airfoil for a gas turbine engine having features for reducing an angle between a cooling aperture and an outer surface of the airfoil to reduce a surface angle of cooling fluid exiting the cooling aperture would be beneficial. Further, an airfoil having a cooling aperture including a change in direction between a first section and a second section of the cooling aperture would be advantageous. Additionally, a method for forming an airfoil for a gas turbine engine, the airfoil having features for improved surface cooling of the airfoil, would be useful.